Carbon-epoxy structures comprise a plurality of carbon fibers embedded in a matrix of resin (typically epoxy) to support and bond the carbon fibers together, and may be provided as pre-impregnated or “prepreg” structures.
Carbon-epoxy structures are strong and light weight, and therefore are used in a wide range of industries either by themselves or co-cured to other carbon-epoxy structures. For example, in the aircraft industry, aircraft wings comprise co-cured carbon-epoxy structures to provide the wings with sufficient strength to handle aerodynamic loads imposed on the wings during flight and sufficiently low weight to accommodate aircraft requirements. Aircraft wings and other aircraft structures are typically formed with arrays of stiffeners or “stringers” co-cured with an outer laminate layer (such as the aircraft skin). Stringers may be fabricated as a prepreg in various cross-sections, such as I-beam, Z, or “hat-shaped” cross-sections. A “hat-shaped stringer” is generally formed by a pair of webs that extend upwardly from a base portion and are connected in a hat shape enclosing a hat-shaped cross-sectional volume.
A lay-up method is typically used to co-cure stringers with the aircraft skin. The aircraft skin, also formed as a prepreg, is laid-up onto the stringer to cover the hat-shaped cross-sectional volume and form a hollow inner volume. Sufficient heat and pressure are applied by autoclave or similar method to co-cure the stringers and the aircraft skin together. A removable mandrel is commonly inserted into the hollow inner volume to support the hollow inner volume during curing. The stringers provide stiffener resistance to loads applied on the aircraft skin, while the hollow inner volume of the stringer provides a light weight in relation to its stiffener strength. The foregoing process may be used for co-curing any type of carbon-epoxy structures, for example, co-curing an aircraft skin to ribs, beams, and other aircraft structural components without compromising the aircraft skin.
An example of a conventional apparatus and method for forming a hollow hat-shaped stringer for aircraft construction and co-curing the stringer to an aircraft skin with an inner removable mandrel for support is illustrated in FIGS. 1-4. As illustrated in FIG. 1, a grooved facesheet 10 has a plurality of grooves 12 spaced at angular intervals and extending in a generally lengthwise or longitudinal direction L. In this example, the grooves 12 have a generally trapezoidal hat-shaped cross-section with corners that curve smoothly and make a radiused transition to a bottom plane 14. Other groove cross-sectional shapes are also possible, such as the semi-circular groove or a V-shaped groove to name a few. The grooved facesheet 10 may be made of a composite material such as carbon fibers or fabric in an epoxy matrix. Other materials may be used singularly or in combination to fabricate the grooved facesheet 10, and other materials may be combined with composites. The materials or combination of materials have a resultant coefficient of thermal expansion CTEhoop of the facesheet (indicated by the double headed arrow in FIG. 1).
FIGS. 2-4 are cross-sectional end views illustrating various stages of a conventional method for forming a stringer 16 co-cured to an aircraft skin using the grooved facesheet 10 of the type shown in FIG. 1. The stringer 16 has a cross-section with a trapezoidal hat-shape having side walls 18 and end flange portions 20a, 20b. Referring to FIG. 2, the stringer 16 is positioned in one of the hat-shaped grooves 12 formed in the grooved facesheet 10. The side walls 18 of the stringer 16 have interior surfaces 22 which define a hollow inner volume 24 within the hat-shaped stringer 16. A rigid, hat-shaped mandrel 26 is positioned within the hollow inner volume 24 of the stringer 16 so that outer surfaces 28 of the hat-shaped mandrel 26 are in supporting contact with the interior surfaces 22 of the stringer 16 to maintain the shape of the stringer 16 under compressive forces applied during lay-up and curing.
Referring to FIG. 3, once the mandrel 26 is positioned inside the side walls 18 forming the hollow inner volume 24 of the stringer 16, an outer substrate layer 30 is laminated over the facesheet 10, stringer end flange portions 20a, 20b, and mandrel 26 to form an aircraft skin. Like the stringer 16, the outer substrate layer 30 may be a prepreg carbon-epoxy structure. The outer substrate layer 30 is co-cured with the stringer 16 by applying sufficient heat and pressure in autoclaving or a similar method. The outer substrate layer 30 becomes bonded by co-curing to the flange portions 20a, 20b of the stringer 16 and the outer surface 28 of the mandrel 26.
Referring to FIG. 4, curing may be performed by first positioning a compressible pad or caul sheet 32 over the outer substrate layer 30, and a vacuum bag 34 is positioned over the caul sheet 32 and the outer substrate layer 30 in order to form a pressurized environment for autoclaving. The air space between the vacuum bag 34 and the outer substrate layer 30 is then evacuated to apply an even pressure against the outer substrate layer 30, end flange portions 20a, 20b of the stringer 16, and outer surface 28 of the mandrel 26. The outer substrate layer 30 is then co-cured to the grooved facesheet 10, the end flange portions 20a, 20b of the stringer 16, and the outer surface 28 of the mandrel 26 at a curing temperature using an autoclave while under vacuum. After curing, the outer substrate layer 30 is separated from the caul sheet 32 and the vacuum bag 34, and the mandrel 26 in the hollow inner volume 24 of the stringer 16 is removed. The outer surface 28 of the mandrel 26 may be made of a material or treated with a material that releases from the interior surfaces 22 of the stringer 16 following curing to permit removal of the mandrel 26.
Any of a variety of mandrels may be used in co-curing of co-cured carbon epoxy structures having a hollow interior volume. For example, there is a hard rubber type of mandrel that may be inserted in the cross-sectional volume of a stringer, but this type of mandrel has proved to be difficult to remove from the stringers after curing. There is also an expendable type of mandrel that may be dissolved after curing, but this type of mandrel requires a messy and impractical dissolution process. Another type is a re-usable elastomeric mandrel system (REMS), which is essentially a rubber mandrel full of ceramic beads. However, the REMS type of mandrel may be impractical to use with longer stringer lengths.
A type of mandrel commonly used is a pressurizable flexible bladder that provides an equal and opposite expansion force so as not to collapse during cure pressures of up to about 100 psi. However, the bladder type of mandrel is not pressurized during lay-up or emplacement, which may allow an outer laminate layer to buckle and cause numerous problems in subsequent operations. Moreover, a bladder type of mandrel may be inconvenient to use because it can cause wrinkles in the outer laminate layer, can have potential leaks which must be checked on every run, must be inserted and removed on every run, and/or replaced at a high cost.
It would therefore be desirable to provide an improved type of mandrel for use in co-curing of hollow carbon-epoxy structures that can reduce or eliminate buckling of an outer laminate layer during lay-up, and that can be removed easily and cleanly after co-curing of the carbon-epoxy structures.